This invention is directed to honeycomb core structures and, in particular, honeycomb core structures that are subjected to bending and shear stresses, such as aircraft control surface structures; and, methods of making such structures.
While the herein described invention was developed for use in the airplane industry by aeronautical engineers and designers, and is described in that environment, it is to be understood that structures formed in accordance with the invention are also useful in other environments. In general, structures formed in accordance with the invention will be useful in many environments requiring bend and/or shear stress resistant panels formed without skin lumps or depressions.
Prior to this invention, solid web, elongated spars were used in many aircraft control surface structures, such as flaps and ailerons to prevent the bending of such structures in a spanwise direction and the shearing of such structures in a chordwise direction. The strength of a solid spar was thought to be required to withstand the bending and shear forces applied to such structures during flight. More specifically, during flight, when airplane control surfaces are in their operative positions, both bending and shear stresses are applied to the control surfaces. These stresses are prevented from damaging the control surface structure by a spar mounted spanwise in the structure. The web of the spar resists both shear stress, which concentrates along the chordwise centerline of the control surface structure, and bending stress, which occurs at right angles to the longitudinal axis of the spar. In essence, therefore, the spar forms a primary load carrying member that resists bend and shear stresses. As noted above, because these stresses are high, in the past, it was thought that a solid web spar was required. Obviously, solid web spars are undesirable because they add unnecessary weight to the control surface structures. Added weight, of course, decreases fuel economy as well as increases the power needed to move the control surfaces. However, the additional weight added by solid spars is not their main disadvantage. The main disadvantage of solid spars is that they cannot be inexpensively produced and still meet exact dimensional requirements. Moreover, solid spars often suffer from web warpage. In the past, these disadvantages have been overcome by bolting or rigidly affixing a spar to a "tooling" platform during the formation of airplane control surface structures. The rigid platform was used to maintain the solid spar in a fixed position (attached to the lower skin) during subsequent forming and bonding steps, hereinafter described.
Turning now to a discussion of the necessity for a solid spar to be precisely formed when used as a strengthening member in a honeycomb core structure (or conversely the necessity that the adjacent honeycomb to be formed in a manner that compensates for spar dimensional variations); it is well known that an adhesive layer located between a skin and a honeycomb core will sink into the porous honeycomb core. As a result, the bonding layer will be thinner than the starting adhesive layer, ignoring any adhesive shrinkage. Conversely, an adhesive layer will not sink or decrease in thickness when applied between a skin and a solid surface, such as the flange of a spar, again ignoring adhesive shrinkage and assuming adhesive is not forced or squeezed out between the solid surfaces. Therefore, when a honeycomb core and a spar flange are joined in a planar manner and a skin applied over the core and the flange, aerodynamically harmful rippling or indentations may occur in the skin surface, if the adhesive used is of uniform thickness. This result can be alleviated by forming the portion of the honeycomb core adjacent the spar such that it has a greater thickness than the spar, i.e. the junction between the core and the spar flange is discontinuous. It is known that this discontinuity should fall between 0.00 and 0.01 inch (average 0.005) if a reliable bond without rippling or identations is to be obtained. In the past, using solid spars, this result has been accomplished using the two stage hot bonding process described below.
In the first stage of a two stage hot bonding process, a lower skin is laid out on the "tooling" platform noted above and an adhesive is applied to the skin surface. The solid spar (usually "C" or I shaped in cross-section) is attached to the skin and bolted in place. Next, the honeycomb core is attached to the web of the spar with an adhesive. At the same time the honeycomb core is attached to the lower skin. Flash tape, protective film and bleeder cloth are placed over the core; and, the structure is sealed in a bag mold and placed in an autoclave to cure the adhesive so that bonds are formed. After the bonds are formed, the protective film, bleeder cloth and the flash tape are removed. In the second stage of this process the exposed honeycomb core is machined to a desired shape. At this time the exposed upper surface of the upper flange of the spar is used as an index point to achieve a 0.005 inch average discontinuity between the upper surface of the spar and the region of the honeycomb core adjacent to the spar. (It is pointed out here that this average discontinuity is extremely difficult to achieve in structures having lengths greater than 10 feet.) After the core is cleaned by a vapor degreasing process, the upper skin is adhesively attached to the core and the spar; and the adhesive is cured in an autoclave so that the upper skin becomes bonded to the core and the spar.
Because of the potential cost savings in man-hours, materials and energy, those skilled in the art have been attempting to find a single stage hot bonding process that can be used to produce reliably bonded aircraft control surface structures, such as flaps and ailerons. One attempted solution ignores the tolerance problem created by the nonuniform dimensions of the spar. In this solution, the honeycomb core was machined such that its outer surface adjacent the spars would be 0.04 inch greater than the outer surface of the spars. The flap or aileron was then assembled in a single stage. During assembly, an extra coating of low flow adhesive was applied on the spars by hand to cover up the mismatching created by the nonuniform dimensions of the spars. Then, the entire assembly was heated in an autoclave to form the adhesive into bond. This attempt to provide a single stage bonding process has a number of disadvantages. Specifically, the use of extra adhesive adds to the weight of the resulting structure. Further, the handwork required to apply the extra adhesive adds manufacturing time and materials and thus increases the cost of the structure. Also, the thick adhesive about the spar area increases the likelihood of leak paths extending to the honeycomb core from the exterior of the structure. Finally, the bond between the skin and the spar, in the area of extra adhesive, has been found to be unreliable.
Therefore, it is an object of this invention to provide new and improved composite structures suitable for withstanding bending and shear forces.
It is also an object of this invention to provide new and improved honeycomb core structures suitable for withstanding bending and shear forces.
It is another object of this invention to provide new and improved aircraft control surface structures, such as flaps and ailerons, that are lighter in weight than similar prior art structures.
It is a further object of this invention to provide composite structures, such as airplane control surface structures, formed by a single stage hot bonding process.
It is yet another object of this invention to provide a new and improved single stage hot bonding process suitable for forming bend and shear stress resistant composite structures, particularly honeycomb core composite structures.